In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot combustion gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine by impingement against an airfoil section of the turbine blades and vanes, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward.
The hotter the combustion and exhaust gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the combustion and exhaust gas temperatures. The maximum temperature of the combustion gases is normally limited by the materials used to fabricate the hot-section components of the engine. These components include the turbine vanes and turbine blades of the gas turbine, upon which the hot combustion gases directly impinge, and other hot-section components as well. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at temperatures of up to about 2150° F. These components are also subject to damage by oxidation and corrosive agents, as well as impact damage and erosion by particles entrained in the combustion gas stream.
Many approaches have been used to increase the operating temperature limits and service lives of the hot-section components such as the turbine blades and vanes to their current levels, while achieving acceptable oxidation, corrosion, erosion, and impact resistance. The composition and processing of the base materials themselves have been improved. Portions of the external surfaces of the turbine blades and vanes are coated with a protective aluminide environmental coating and, in some cases, an insulating thermal barrier coating is applied overlying the aluminide coating.
Physical cooling techniques are also used. In one practice, the component is provided with an internal cooling cavity. Compressor bleed air flows into the internal cooling cavity through an inlet in the root of the component, and out of the internal cooling cavity through small-diameter outlets, such as found at the leading edge, trailing edge, and faces of the turbine blade. The air flow carries heat away from the cooled component and also provides film cooling. This physical cooling technique has allowed the hot-section components to run even hotter than possible for uncooled components.
As the operating temperatures have been pushed even further upwardly, it has been found that the internal surfaces of the internal cooling cavities experience oxidation and other degradation during service, particularly near the tip of the turbine blade. Therefore, more recently internal coatings have been applied to protect the internal surfaces of the hollow components.
However, the application of coatings to the internal surfaces poses some problems not found for external coatings, particularly because of the presence of the small-diameter outlets. Existing external coating techniques, when applied to internal coating, and candidate internal coating techniques are not sufficient to achieve the necessary internal coatings.
There is a need for an improved approach to coating internal surfaces of hollow articles having small-diameter outlets. The present invention fulfills this need, and further provides related advantages.